The invention relates to supplying electrical power to an aircraft, and more particularly to electrical equipment of an aircraft engine and/or of its environment.
The field of application of the invention is more specifically that of airplane engines, in particular gas turbine engines. Nevertheless, the invention is also applicable to helicopter engines.
The term “electrical equipment of an aircraft engine or of its environment” is used herein to mean not only electrical equipment useful for the actual operation of the engine, but also electrical equipment associated with the engine pod, such as, for example: electrical circuits for de-icing, or electromechanical actuators for reversing thrust on gas turbine airplane engines; or even equipment associated with the wing carrying the engine, such as, for example: de-icing or anti-icing electrical circuits for the airplane wing.
A traditional layout for producing and distributing electricity from a gas turbine airplane engine is shown in FIG. 1.
Two generators 1, 1′ (or more than two for redundancy purposes or for optimizing the generation of electrical power depending on the application in question) are mounted on an accessory gearbox (AGB) that is mechanically coupled to a turbine shaft of the engine. The generators are typically starter/generators (S/Gs) comprising a synchronous generator that is associated with an exciter and that supplies an alternating voltage at a frequency that varies as a function of the speed of the engine, the assembly comprising the exciter and the synchronous generator being controlled to operate in synchronous motor mode when starting the turbine.
The alternating voltages supplied by the generators 1, 1′ are conveyed by lines 2, 21 to an electrical power distribution network 3 on board the airplane, referred to as the “on-board network”. A circuit 4 of the on-board network connected to the lines 2, 2′ supplies a regulated alternating voltage typically of 115 volts AC (Vac) or 230 Vac, on one or more distribution buses. The circuit 4 also powers a voltage converter 5 that delivers a regulated direct current (DC) voltage, typically 270 Vdc or ±270 Vdc, on one or more buses. The voltages supplied by the circuits 4 and 5 are fed to the various electrical loads on board the airplane, mainly in the fuselage zone.
In the engine, an electronic engine control unit 6 (ECU) is powered by a generator 7 such as a permanent magnet alternator (PMA) mounted on the accessory gearbox (AGB). The ECU is also connected to one of the buses 4, 5, e.g. to the regulated alternating voltage bus 4 in order to be powered properly so long as the speed of the engine is not sufficiently fast to enable the PMA to supply the required electrical power, or in the event of the PMA failing. The ECU uses the electricity it receives to enable its components to operate and to excite the various elements of the engine that require limited amounts of electrical power such as probes or sensors, actuators, or servo-valves.
There is a present trend to replace hydraulic power more and more with electrical power for actuating various pieces of equipment in an aircraft engine or its environment. Thus, some airplanes are fitted with thrust reversers that are electrically actuated, such that an electrical power supply line 8 must connect the on-board network 3 of the airplane to such an electric thrust reverser 9. Such a line is in addition to those needed for powering static equipment, such as lines 10, 11 for powering de-icing circuits 12, 13 of the engine pod and of the wing carrying the engine.
Conveying electricity from the on-board network to loads outside the fuselage by means of lines that must be carefully secured and insulated represents a considerable amount of weight and bulk, running the risk of becoming ruling dimensions, or even of being excessive if the amount of equipment to be powered increases.